A turbulence generating grid was designed and installed in the turbine inlet which produced the target nominal value of 10 percent free stream turbulence. Aerodynamic documentation of the rotor and stator midspan surface pressure distributions were obtained. Midspan heat transfer data were obtained on the rotor and stator for variations in inlet turbulence, rotor-stator axial spacing, and rotor incidence

Blair, Michael F.

Dring, Robert P.

Joslyn, H. David

English

NASA Technical Reports Server

MEASUREMENT OF AIRFOIL HEAT TRANSFER
COEFFICIENTS ON A TURBINE STAGE*
N88- 11158
Robert P. Dring
Michael F. Blair
H. David Joslyn
United Technologies Research Center
East Hartford, Connecticut 06108
INTRODUCTION
The primary basis for heat transfer analysis of turbine airfoils is
experimental data obtained in linear cascades. These data have been very
valuable in identifying the major heat transfer and fluid flow features of
a turbine airfoil. The question of major interest is how well all of these
data translate to the rotating turbine stage. It is known from the work of
Lokay and Trushin (Ref. I) that average heat transfer coefficients on the
rotor may be as much as 40 percent above the values measured on the same
blades non-rotating. Recent work by Dunn and Holt (Ref. 2) supports the
conclusion of Ref. I. What is lacking is a set of data from a rotating
system which is of sufficient detail as to make careful local comparisons
between static cascade and rotor blade heat transfer. In addition, data is
needed in a rotating system in which there is sufficient documentation of
the flow field to support the computer analyses being developed today.
Other important questions include the impact of both random and periodic
unsteadiness on both the rotor and stator airfoil heat transfer. The
random unsteadiness arises from stage inlet turbulence and wake generated
turbulence and the periodic unsteadiness arises from blade passing effects.
A final question is the influence, if any, of the first stator row and
first stator inlet turbulence on the heat transfer of the second stator row
after the flow has been passed through the rotor.
OBJECTIVES
The first program objective is to obtain a detailed set of heat
transfer coefficients along the midspan of a stator and a rotor in a
rotating turbine stage (Fig. I). These data are to be such that the
rotor data can be compared directly with data taken in a static cascade.
The data are to be compared to some standard analysis of blade boundary
layer heat transfer which is in use today. In addition to providing this
all-important comparison between rotating and stationary data, this
experiment should provide important insight to the more elaborate full
three-dimensional programs being proposed for future research. A second
program objective is to obtain a detailed set of heat transfer coefficients
along the midspan of a stator located in the wake of an upstream turbine
stage. Particular focus here is on the relative circmnferential location
of the first and second stators. Both program objectives will be carried
out at two levels of inlet turbulence. The low level will be on the order
of 1 percent while the high level will be on the order of I0 percent which
is more typical of combustor exit turbulence intensity. The final program
*Work done under NASA Contract NAS3-23717.
165
https://ntrs.nasa.gov/search.jsp?R=19880001776 2020-03-20T08:34:19+00:00Z
objective is to improve the analytical capability to predict the
experimental data.
PROGRESS
During the past year the following phases of the progrnm were
completed: (I) A turbulence generating grid was designed and installed in
the turbine inlet which produced the target nominal value of 10Z
£ree-strenm turbulence. (2) AerodynJmic documentation of the rotor and
ststor aidspan surface pressure distributions was obtained. (3) Midspan
heat transfer data were obtained on the rotor and stator for variations in
inlet turbulence, rotor-stator axial spacing, and rotor incidence. Each of
these three areas will now be discussed in greater detail.
I. Inlet Turbulence
As part of the present contract the effects of high levels of
free-stregm turbulence on the heat transfer distributions through the LSRR
turbine blading will be examined. These heat transfer data will be
obtained for a turbine inlet turbulence intensity of approximately 10
percent. This is typical o£ the level of turbulence measured at the exit
of aircraft gas turbine combustors. For purposes of airfoil to airfoil
consistency (circumferential uniformity) and so that the present results
can be compared with other data on the effects of turbulence on heat
transfer, the turbulence generated for these tests is required to be
spatially uniform, nearly isotropic, and temporally steady (over time
scales long when compared to the turbulent fluctuations). In addition the
test turbulence must be generated in a manner such that there i8 a
reasonably high intensity through the 1 1/2 stages of the turbine, i.e.,
the stres_wise decay of the turbulence must be similar to that in an
engine.
The turbulence generator selected consisted of a nearly square array
lattice of three concentric rings spaced uniformly in the radial direction
with 80 radial bars evenly spaced circumferentially. Both the rings and
radial bars are of nearly square 1/2 inch cross-section. The mesh spacing
of the bar is 2.1 inches radially and 4.5 degrees (2.1 in. at mid-annulus)
circumferentially.
Prel/minary indications are that without the grid installed the inlet
turbulence was approximetely 0.§X at an axial location 22Z of axial chord
ahead o£ the first stator leading edge. With the grid installedp at this
s/me axial location, the inlet turbulence intensity was typically 9.8Z.
The spanwise distributions at four different circumferential locations
(relative to the stator leading edge) are shown in Fig. 2. The data
indicate that the turbulence is spatially uniform_ nearly i8otropic, and
temporally steady.
2. Aerodynamics
The aerodynamic documentation of the turbine stage indicated that all
parameter8 were very c_ose to data obtained during prior testing with this
turbine model. As an exsnple, the stator and rotor pressure distributions
are shown in Figs. 3a and 3b for the case with the small (15Z) axial gap,
design flow coefficient (Cx/Um=0.78), and the inlet turbulence gener-
166
ating grid installed. Agreement with a potential two dimensional flow
calculation at this midspan location is excellent. The computed surface
velocity distributions are used as the input to the suction and pressure
surface boundary layer calculations.
3. Heat Transfer
The nethod of fabrication of the heat transfer models was described in
detail in the 1984 HOST review. Suffice it to say here that Joulean
heating of a thin sheet of metal foil on the stator and rotor surfaces
produces a nearly uniform surface heat flux. Conduction and radiation
effects produce small departures from complete uniformity. Local airfoil
surface tesaperature8 are measured using thermocouples welded to the back of
the foil and the air tezperature i8 measured using thermocouples in the air
stremn. The secondary junctions to copper wire are all made on Uniform
Temperature Reference blocks (Keys Instruments, UTR-ASN) and the data were
recorded using a Hevlett-Packsrd 300 channel data aquisition unit
(3A97A/3A98A), and an ice point reference (Keys Instruments, KI&0-A). A
212 ring slip-ring unit (Wenden Co.) was used to bring heater power onto
the rotor and to bring out the thermocouple data.
A low speed unheated test run indicated a max-toIL_n variation in
absolute tenperature for all thermocouple8 in all airfoils (rotor and
8tator, suction and pressure surfaces) of _ 0.3°F. As part of the
procedure of taking the model data the temperature of the flow just
upstrem of the grid location was monitored at eight equally spaced
circumferencial locations. For these initial test cases it was decided
that data would be recorded only if the outside mbient (rig inlet)
conditions were such that the inlet temperature was below 65°F and that
the tmaperature range of the eight inlet thermocouples was within
approximately _ I°F. These criteria were intended to assure that there
would be both an adequate temperature difference between the heated
surfaces and the freestre_m and a uniform temperature approach flow for the
stator and rotor models. Preliminary examination of these data indicates
that there was excellent run-to-run and day-to-day repeatability of results
for nominally identical test conditions.
A sample set of rotor heat transfer distributions is presented in
Figure A. These data were all obtained with an axial rotor-stator
separation of 15Z chord. Data are presented for cases both with and
without the turbulence grid installed for three flow coefficients
(C /U.). Although the rotor inlet relative velocity (W 1) differed
wi_el_ for these three flow coefficients (WI=91.A , 115.0 and 163.3 ft/sec
for C_/U_-0.68_ 0.78 and 0.96 respectively) the rotor exit relative
velocity'was invariant at W2=176 ft/sec. For ease of comparison the
Stanton numbers of Figure A-were computed using the nearly constant rotor
exit conditions. Computation of predicted heat transfer distributions for
these test cases was not completed at the time of submittal of this interim
report so the data are simply connected with straight line secants.
167
The three sets of data in the lower portion of Figure 4 (open symbols)
were obtained without the turbulence generating grid installed. A
comparison of the relative levels of Stanton numbers for these three cases
indicates that: (1) the heat transfer rates near the trailing edge of both
the suction and pressure surfaces are independent of flow coefficient and
(2) the peak stagnation region (S=0) heat transfer obeyed the classic
cylinder leading edge correlation with St inversely proportional to the
square root of the Reynolds number based on nose diameter and approach
relative velocity. The most distinct difference between the three
low turbulence heat transfer distributions resulted near S _1 inch. As
C_/U_ increased, a region of very high local Stanton number resulted at
t_xs location. The basic cause of this spike in heat transfer i8 the
suction surface overspeed (S ~3/4 in.), the strength of which is a function
of the flow coefficient. At Cx/Um=0.68 there i8 only a slight
overspeed followed by a favoraSle-pressure gradient to midchord while for
C_/U_=0.96 the suction surface overspeed location has by far the
h_gh_st velocity on the airfoil and is followed by i very strong adverse
pressure gradient. For C/Um=0.96 the boundary layer is apparently
unable to negotiate the aaverse pressure gradient, separates, passes
through a very short transition and reattaches as a high-speed,
fully-turbulent layer. For C /U_=0.68 the boundary layer seems toX m
experience an extended transition length through the favorable pressure
gradient to near midchord. A much subdued version of this same phenomenon
can be seen on the pressure surface where the most severe overspeed
(S--1/2 inch) occurs for Cx/Um=0.68.
A general increase in the levels of heat transfer was Observed for the
three cases with enhanced free-stream turbulence (solid symbols). Two of
the most extreme local effects were: (I) near the suction surface overspeed
where the freestream turbulence has severely increased the heat transfer
and (2) at S ~-2 inches for C_/U_=0.96 where the free-stream turbulence
produced a much shorter transltlon length.
0ff-uidspan data (not shown here) over the aft portion of the suction
surface indicate much higher Stanton numbers near the root and tip. These
enhanced heat transfer levels were almost certainly a product of the
passage and tip-leakage vortex systems. It is anticipated that the flow
convergence associated with these sane corner vortices viii cause the
measured mid-span trailing edge Stanton numbers to fall well below values
predicted for 2-d flow.
REFERENCES
Lokay, V. I., and Trushin, V. A.: Heat Transfer from the Gas and
Flow-Passage Elements of a Rotating Gas Turbine. Heat Transfer -
Soviet Research, Vol. 2., No. 4, July 1970.
. Dunn, M.G., and Holt, J.L.: The Turbine Stage Heat Flux
Measurements. Paper No. 82-1289, AIAA/ASME 18th Joint Propulsion
Conference, 21-23, June, 1982, Cleveland, Ohio.
168
TURBINE STAGE AT 15% AXIAL GAP
Rotation
Stator 1
Rotor
FIG. 1
169'
STREAMWISE TURBULENCE (RMS)
14
12
8
Center 60_ of span
_ Mean 9.8_
A A
- i- _ __ __ ,,"
B
6 I I I I
0 120 40 60 80
Span
100
Figure 2
170
oRIGINAE PAGE Ig
OF POOR QUALITY
FIRST STATOR
PRESSURE DISTRIBUTION
LOItR06 • FZR|T VRN[ I Im.O Z I_Pi_l • Ili4Y*O| IIm2200|O1|O4
PtELZRZ_Y NEON |OI2200|5_0|0
iUN 0i, POZNT 0, CX_ * i.,0i, 0JZD ZN, 8_P m |1_ |H220B|S|SO4
IEMRLYSZ0 MSH 1f1220S|SIII4
IEAN_LYSZS $1R22i5152034
PLOT 0r CP (R_ZNPP). 0rTI0 ; 90.N01 |[T20 • |10.0081CP • (P-P2 )/(P0|-P2) _| O,0001 R2 * O,00Ol
i0RoRL[/MI• 0.0000 OPT/G1 • 0.2224 G_RR_ • 1.400O
ALF|0 • 90.1iS0 PZTCH/0X * i.2999 H2/N! • 0.9957
I,OO00 .
O.O000 .
0.0000 .
0.40e0.
0.2000--
O.OOO0
-0.2000 --
4.41_0 --
-O.$OO0 _
-4.0000 --
-8.0OOO
I I I I I I I I I I
O.OOO0 • ION O.IO00 O.IOO0 0.4000 O.O000 O.OOO0 0.,000 O.OOOe 0.9000 | OO0
Figure 3a
171
ORIGII_AI3 P_iGE IS
OF POOR QUALITY
ROTOR PRESSURE DISTRIBUTION
LSRR?S. iOSOO / F_tST 0LRDE / 041.t/ 0PAN :0122t010110_
P•ELZRZNRIY mESH 111112100151250
IUU 10o POZmt l° exam - I.PII, OtZO ZN. GflP• 18_ 10'132110191111
tlmmmLyszs mtsm 1H2215155120
NERNRLYSZS |lHI220515)PO5
PL0? OF CP (|RSECN). IETID • 40.0001 IlY2O " 155.0000
CP • (P-P2 )/(PSI-P2) Ri • i*ilOO M2 • 0,0000
OIq*RLE/MI8 0.leOl Df'T/01 • l. llll CI41_qA • 1.4005
ALFID • 14O.eOO0 PXTCN/|M • 0.9557 N2/N| • 1,0141
1.0110 _
l. OlO0 _
0.0000 --
0.4015 --
0.2000 _
O.OI0O
-0,2100 _
• 4111
0
t 0000 t
00.00•0
= oo,, I I I I I I I I I I
* .
0.OO00 •.10Ol 1.2008 1.10N 0.4000 O.•000 O.•000 0,?0N 1.0000 0.9000 l.•OOl
Figure 3b
172
ROTOR HEAT TRANSFER
DISTRIBUTIONS
Close rotor.stator spacing
0.01
0.008
0.006
0.004
0.002
]°-
.
II
0
0
0
0 I
-8
' I ' I ' I ' 1
Open symbols -- grid out (low turbulence)
Solid symbols -- grid in (high turbulence)
CxlU = 0.960
0.78
0.68
0.96
0.78
0.68
Pressure surface } Suction surface
I
= I I I , i I I,
-4 0 4
S _ INCHES
I I
8
i
12
Figure 4
173


Measurement of airfoil heat transfer coefficients on a turbine stage

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